In a gas turbine engine, air is pressurized in a compressor and is mixed with fuel in a combustor for generating hot combustion gases which flow downstream through turbine stages that extract energy therefrom. The turbine includes stationary airfoils (vanes) that direct the combustion gases through respective downstream rows of rotating airfoils (blades) extending radially outwardly from a rotating shaft.
Present-day high performance turbines include vanes that are capable of withstanding temperatures approaching 1600° C. or higher. While high temperature metal alloys and ceramic materials may be used for constructing the vanes and blades, active cooling of the structures with a cooling fluid is required in many applications. Cooling is typically accomplished by directing cooling air through the hollow cavity of the airfoil.
Various schemes have been used in the past to actively cool gas turbine components such as the stationary vanes. For example, in U.S. Pat. No. 5,772,398, entitled COOLED TURBINE GUIDE VANE, a cooled turbine vane is disclosed as including a hollow aerodynamic portion between inner and outer platforms. The interior of the aerodynamic portion is partitioned into a leading edge duct and a main cavity in which a perforated tubular member is disposed, being spaced from the interior and exterior side walls of the vane by longitudinal ribs. The tubular member is divided by a partition into two cavities on the interior and exterior sides of the partition. A first cooling circuit includes the leading edge duct and the interior cavity of the tubular member, and a second cooling circuit includes the exterior cavity and a cooling system for the inner platform, both circuits being supplied with cooling air by the same source from the outer platform. Cooling air from each circuit passes through the perforations of the tubular member to impinge on the inside face of the respective side wall of the vane and is then guided toward the trailing edge, where it escapes through slits in the trailing edge wall.
Another example of a prior art device is disclosed in U.S. Pat. No. 5,813,827, entitled APPARATUS FOR COOLING A GAS TURBINE AIRFOIL. This apparatus includes two radially extending passages connected to the outer shroud to direct a cooling fluid to a plenum formed about mid-span adjacent to the trailing edge. Two arrays of cooling fluid passages extend from the plenum. One array extends radially inward toward the inner shroud. The plenum distributes the cooling fluid to the two arrays of passages so that it flows radially inward and outward to manifolds formed in the inner and outer shrouds. The manifolds direct the spent cooling fluid to a discharge passage.
To utilize the cooling air passing through a gas turbine vane effectively, it is useful to reduce the size of the cooling passage, since cooling air traveling along the center of the passage is not in contact with the surface being cooled. However, the reduction of the internal cooling passage cross-sectional area to a desired degree for cooling purposes would result in an undesirably thin aerodynamic shape or the necessity for installing complicated and costly structures within the vane for directing the fluid flow. A thinner aerodynamic shape requires a larger number of airfoils to produce the desired aero performance, thereby increasing cost and reducing efficiency. A means of improving cooling efficiency without affecting the external airfoil contour is desired.